Impingement cooling for turbofan exhaust assembly

ABSTRACT

A system comprises a bypass duct, a turbine case, a turbine exhaust ring and a splash plate. The turbine case is formed along a radially inner margin of the bypass duct, and the turbine exhaust ring is coaxially disposed within the turbine case. The splash plate extends axially along the turbine exhaust ring, radially spaced between the turbine exhaust ring and the turbine case. There are cooling fluid apertures in the turbine case to provide cooling fluid flow from the bypass duct onto the splash plate, and impingement holes in the splash plate to provide impingement flow onto the turbine exhaust ring.

CROSS-REFERENCE TO RELATED APPLICATION

This application is a continuation of Snyder et al., IMPINGEMENT COOLINGFOR TURBOFAN EXHAUST ASSEMBLY, U.S. patent application Ser. No.12/167,780, filed Jul. 3, 2008.

STATEMENT OF GOVERNMENT INTEREST

This invention was made with U.S. Government support under Contract No.N00019-02-C-3003, awarded by U.S. Navy. The U.S. Government may havecertain rights in the invention.

BACKGROUND OF THE INVENTION

This invention relates generally to gas turbine engines, andspecifically to cooling techniques for a turbine exhaust assembly. Inparticular, the invention concerns impingement cooling of a low-bypassturbofan exhaust case (TEC) assembly, with applications in militaryaviation and high-performance aircraft.

Standard gas turbine engines are built around a power core comprisingcompressor, combustor and turbine sections, which are arranged in flowseries with an upstream inlet and a downstream exhaust nozzle. Thecompressor compresses air from the inlet. The compressed air is used anoxidant in the combustor, and, in some applications, for accessorypneumatic functions and environmental control. Fuel is injected into thecombustor, where it combines with the compressed air and ignites toproduce hot combustion gases. The hot combustion gases drive the engineby expansion in the turbine section, and are exhausted to the nozzlethrough a turbine exhaust case (TEC) assembly.

The turbine section drives the compressor via a rotating shaft, or, inmost larger-scale applications, via a number of coaxially nested shaftsand independently rotating turbine/compressor assemblies or spools. Eachspool, in turn, employs a number of stages, in which rotating bladescoupled to the shaft are alternated with stationary vanes coupled to ashroud or other fixed component of the engine housing.

Energy that is not used to drive the compressor and accessory functionsis available for extraction and use. In ground-based applications,energy is typically delivered in the form of rotational motion, which isused to drive an electrical generator or other mechanical load coupledto the shaft. In aviation applications, the gas turbine engine alsoprovides reactive thrust.

The relative contributions of rotation and thrust depend upon enginedesign. In turbojet engines, for example, which are an older design,essentially all the net thrust is generated in the exhaust. In modernturbofan engines, on the other hand, the shaft is used to drive a ductedpropeller or forward fan, which generates additional thrust by forcingair through a bypass flow duct surrounding the engine core.

Turbofan engines include low-bypass turbofans, in which the bypass flowis relatively small with respect to the core flow, and high-bypassturbofans, in which the bypass flow is greater. High-bypass turbofanstend to be quieter, cooler and more energy efficient, particularly insubsonic flight applications for commercial and other general-purposeaircraft. Low-bypass turbofans can be somewhat louder and less fuelefficient, but provide greater specific thrust. For these and otherreasons, low-bypass turbofans are generally utilized in military jetfighters and other high-performance supersonic aircraft.

In supersonic applications, the turbofan engine is typically providedwith an afterburner. Afterburning systems provide thrust augmentation byinjecting additional fuel into an augmentor assembly, downstream of theTEC, where it mixes with the core flow and ignites to increase thethrust. Afterburning substantially enhances engine performance, but isalso associated with additional costs in efficiency, noise output andthermal signature.

The main design goals for aviation-based gas turbine engines areperformance, efficiency, reliability and service life. Performance andefficiency both favor higher combustion temperatures, which increase theengine's specific thrust and overall thermodynamic efficiency.Unfortunately, higher combustion temperatures also result in increasedthermal and mechanical loads, particularly for engine components alongthe hot gas flowpath, downstream of the combustor. This can affectservice life and reliability, and increase operational costs associatedwith maintenance and part replacement.

In high-performance (low-bypass) turbofans, gas path temperatures areoften a factor at the TEC assembly, where hot combustion gases flow fromturbine section (upstream of the TEC) toward the afterburner/augmentor(downstream of the TEC). The issue can be problematic proximate theforward outer diameter ring (FODR), on the upstream end of the TECassembly. In this region, operational conditions can sometimes establisha negative pressure differential between the FODR plenum, whichsurrounds the FODR, and the exhaust gas flow, inside the FODR.

Negative FODR plenum overpressure allows hot gas inflow, impairingcooling efficiency. The result is decreased service life and increasedrisk of mechanical failure. There is thus a need for improved TECassembly cooling techniques that provide increased service life andreliability without sacrificing performance and efficiency.

BRIEF SUMMARY OF THE INVENTION

This invention concerns a turbine system, for example a turbine exhaustassembly or cooling system therefore. The system includes a bypass duct,a turbine case, a flow ring and a splash plate.

The bypass duct has a radially inner margin and a radially outer margin,with the turbine case formed along the radially inner margin. The flowring is coaxially disposed within the turbine case, and the splash plateextends axially along the flow ring, radially spaced between the turbineexhaust ring and the turbine case.

A plurality of cooling fluid apertures are formed in the turbine case,in order to provide cooling fluid flow from the bypass duct onto thesplash plate. A plurality of impingement holes are formed in the splashplate, in order to provide impingement flow onto the flow ring.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a cross-sectional schematic of a turbine exhaust case (TEC)assembly, located in an aft portion of a turbofan engine.

FIG. 2 is an expanded cross-sectional schematic of the TEC assembly inFIG. 1, showing high-pressure cooling fluid supply apertures, a splashplate and impingement holes.

FIG. 3 is an additional cross-sectional schematic of the TEC assembly inFIG. 2, showing film cooling openings.

DETAILED DESCRIPTION

FIG. 1 is a cross-sectional schematic of turbine/turbofan exhaust case(TEC) assembly 10, located in an aft portion of turbofan engine 11. TECassembly 10 and turbofan 11 form generally annular or cylindrical bodiesabout axial centerline C_(L), with FIG. 1 showing the top cross sectiononly. The illustrated region of turbofan 11 is aft (downstream) of thecompressor and combustor sections (not shown in FIG. 1), and forward(upstream) of the exhaust nozzle (also not shown).

TEC (turbine exhaust case) assembly 10 comprises forward outer diameterring (FODR) 12, finger seal 13, forward inner diameter ring (FIDR) 14and TEC leading edge piece 15. TEC leading edge piece 15 comprises acombination of strut or support rod components to space FODR 12 fromFIDR 14, which are contained within fairing assembly 16. In someembodiments, TEC assembly 10 also includes one or more elements of thelow-pressure turbine (LPT) section, such as final LPT stator (vane)stage 17.

FODR 12, finger seal 13, FIDR 14, TEC leading edge piece 15 and theother components of TEC assembly 10 are typically manufactured fromdurable heat-resistant materials such as high-temperature metal alloysor superalloys. This protects the components from extreme operationalconditions due to the flow of high-temperature exhaust (hot combustiongases), downstream of the combustor and turbine sections. In someembodiments, one or more of these components is provided with a thermalor protective coating, such as a ceramic thermal barrier coating (TBC),an aluminide coating, a metal oxide coating, a metal alloy coating, asuperalloy coating, or a combination thereof.

TEC assembly 10 forms a generally annular flowpath or flow region forcore working fluid (exhaust gas) flow between LPT stage 17 andafterburner assembly (augmentor) 18. FODR 12 forms an outer flow ringaround the annular exhaust flowpath, and FIDR 14 forms an inner flowring inside the annular exhaust gas flowpath. Thus FODR 12 defines aradially outer boundary for the combustion gas flow, and FIDR 14 definesa radially inner boundary for the combustion gas flow.

Forward outer diameter ring plenum (FODR plenum) 19 forms a generallyannular cavity, coaxially oriented about the hot exhaust gas flowpathinside FODR 12, and within the relatively cool bypass flowpath in bypassduct 23, described below. More specifically, FODR plenum 19 comprises aplenum section with a radially inner margin (or boundary) defined atleast in part by FODR 12, along the hot exhaust gas flowpath, and aradially outer margin (or boundary) defined at least in part by LPT case20 or inner case 21, or both, along the cool bypass flowpath.

LPT case (or turbine case) 20 is coaxially disposed about the turbinesection of turbofan 11, which extends upstream of TEC assembly 10. Innercase (or exhaust case) 21 extends downstream of LPT case 20, and iscoaxially disposed about TEC assembly 10.

Finger seal 13 forms a pressure and fluid flow seal between FODR 12 andLPT case 20 at the forward (upstream) end of FODR plenum 19, and FODR 12forms a pressure and fluid flow seal with inner case 21 at the aft endof FODR plenum 19. This defines the axial extent of FODR plenum 19, fromthe forward end of FODR 12 proximate finger seal 13, to the aft end ofFODR 12 downstream of probe assembly 22.

Hot exhaust gases flow downstream through turbine exhaust case 10 fromfinal turbine stage(s) 17 toward augmentor 18. The core exhaust gasflowpath comprises a region of divergent and radial/axial flow F, in theforward or upstream section of TEC assembly 10, and a region ofsubstantially axial flow F′, in the aft or downstream section of TEC 10.

Probe assembly 22 extends from FODR plenum 19 through FODR 12, into thecore exhaust gas flowpath between FODR 12 and FIDR 14. Probe assembly 22is used to measure hot gas path fluid parameters, as described in moredetail with respect to FIG. 2.

Bypass flow duct 23 forms an annular channel for bypass flow around theengine core. In the region of TEC assembly 10, the inner margin ofbypass flow duct 23 is defined or formed along LPT case 20 and innercase 21, and coincides with the radially outer margin of FODR plenum 19.The outer margin of bypass duct 23 is defined along outer engine casing(or fan cowling) 24. Bypass duct 23 also extends downstream of TECassembly 10, past blocker door 25 toward afterburner/augmentor assembly18.

Cooling fluid plenum 26 is formed about the aft end of TEC assembly 10,between inner case 21 and bypass flow duct 23. In some embodiments,cooling fluid plenum 26 extends downstream of TEC assembly 10 towardaugmentor 18. In these embodiments, cooling fluid plenum 26 typicallyextends to modulated exhaust cooling (MEC) module 27, which providesflow during afterburning operation of augmentor 18.

Blockers 25 are typically formed as vanes or fixed blades with anairfoil cross section. Each blocker (or blocker door) 25 comprisesforward (upstream) blocker section 28, aft (downstream) blocker section29 and blocker seal 30. Blocker seal 30 is typically formed as a fingerseal or similar structure that forms a pressure and flow seal betweenblocker door 25 and a radially inner boundary or margin of bypass duct23, such as inner case 21.

A number of blockers 25 are circumferentially arranged about bypass duct23. Each blocker 25 extends from the radially inner boundary of bypassflow duct 23 to a radially outer boundary or margin of bypass flow duct23, such as engine casing 24. Forward blocker section (or blocker door)28 with blocker seal 30 is configured to “swing” or move in acircumferential sense with respect to aft blocker section 29, reducingthe flow space between adjacent blockers 25.

Blockers 25 form a modulating exhaust cooling device configured forvariable closure and commensurate bypass flow reduction. In particular,blockers 25 function as a variable area device that regulates therelative pressure differential between upstream bypass flow A anddownstream bypass flow A′.

When forward blocker section 28 is aligned with aft blocker section 29,blockers 25 are substantially axially oriented and the bypass flow areais relatively large. In this configuration, bypass flow is increased andthe relative pressure differential between upstream bypass flow A anddownstream bypass flow A′ is decreased. When forward blocker section 28swings or rotates with respect to downstream blocker section 29, suchthat the two sections of blocker 25 are out of alignment, the flow areais reduced and the pressure differential is increased.

Hot gas flow through TEC assembly 10 creates a need for active cooling,particularly in the region proximate FODR 12 and the other upstreamcomponents of TEC assembly 10. To provide efficient cooling and preventinflow in this region, a positive overpressure is desired between FODRplenum 21 and the hot flowpath. In particular, it is desired that thecooling fluid pressure exceeds the combustion gas pressure along thecomplete axial extent of FODR plenum 19 and FODR 12.

Achieving this positive overpressure is complicated by the fact that thehot gas flowpath along FODR 12 tends to exhibit increased pressure inthe downstream (aft) direction. In particular, the radial flow componentof incoming flow F is reduced by TEC assembly 10, such that outgoingflow F′ is straightened and becomes more axial. As the flow straightensit also diverges and slows, as compared between the forward and aft endsof FODR 12, FIDR 14 and FODR plenum 19.

Flow straightening has a number of benefits, including reducedturbulence, improved engine and augmentor efficiency, and lower radarand infrared (thermal) profile. Unfortunately, flow straightening isalso associated with an increase in the hot gas path pressure along theTEC assembly, due to Bernoulli's principle and other effects such asflow stagnation, recirculation and turbulence. In previous enginedesigns, this resulted in a region of negative overpressure along atleast at the aft end of the FODR plenum, particularly during power lifttesting, STOVL (short takeoff or vertical landing) and otherhigh-pressure, high-temperature core flow conditions.

To address this situation, FODR plenum 19 is provided with direct-flowor “main jet” cooling fluid apertures 31. Main jet apertures 31 providedirect cooling flow into FODR plenum 19, utilizing high-pressure flow Afrom the upstream region of bypass duct 23. In particular, main jetapertures 31 are located forward of blocker door 25, where upstream flowA exhibits a relatively higher pressure than downstream flow A′, aft ofblocker door 25.

This distinguishes from designs that utilize lower-pressure coolingfluid sources such as plenum 26, a bleed air system or lower-pressurebypass flow A′, in the downstream region of bypass duct 23. Theseindirect sources provide lower-pressure cooling fluid to FODR plenum 19than direct-flow apertures 31, because they require longer, moreconvoluted and circuitous flow paths, or are otherwise subject toincreased blockages, obstructions, constrictions, friction andassociated loss mechanisms. Indirect sources such as plenum 26 can alsoproblematic when in flow communication with additional downstream loads,such as MEC module 23 or a TEC OD (outside diameter) feed system.

Main jet apertures 31, in contrast, provide sufficient direct coolingfluid flow to generate a positive overpressure along the full axialextent of FODR plenum 19. Adjustable blocker doors 25, moreover, locateddownstream of main jet apertures 31, allow the positive overpressure tobe maintained over a wide range of turbofan operating conditions. Thispositive overpressure enhances TEC assembly cooling, but also implicatesa number technical design challenges, as described below.

FIG. 2 is a cross-sectional schematic of TEC assembly 10, showinghigh-pressure cooling fluid (main jet) apertures 31 and splash plate 32with impingement holes 33. TEC assembly 10 comprises FODR 12, fingerseal 13 and TEC leading edge piece 15, as described with respect to FIG.1, above. Apertures 31 feed high-pressure cooling fluid from bypass duct23 to FODR plenum 19, and splash plate/baffle 32 provides impingementcooling of FODR 12 via impingement holes 33.

As shown in FIG. 2, TEC assembly 10 is configured for use in alow-bypass, high-performance afterburning turbofan. In otherembodiments, the particular details of TEC assembly 10 vary,particularly the size and configuration of FODR plenum 19 and itsrelated engine components, including FODR 12, finger seal 13, TECleading edge piece 15, LPT case 20, inner case 21, probe assembly 22,splash plate 32, and FIDR 14 (which is shown in FIG. 1).

Direct-flow, high-pressure cooling fluid apertures 31 are formed asholes, slots or other openings in one or both of LPT case 20 and innercase 21. Typically, apertures 31 are formed by mechanical drilling,electron beam drilling, laser drilling, laser percussion drilling,electron discharge machining, or another drilling or machiningtechnique. Alternatively, apertures 31 are formed by any combination ofdrilling, machining, stamping, casting and molding, and analogoustechniques.

In some embodiments, apertures 31 are formed only in LPT case 20. Inthese embodiments, apertures 31 are sometimes formed as a single row ofat least ten (10) holes, with a total area of at least two and one halfsquare inches (2.5 in², or about 1,600 mm²). In one such embodiment,apertures 31 comprise approximately eighteen (18) holes, with a diameterof about one half inch (0.50″, or about 12.7 mm) each, and a total areaof about three and a half square inches (3.5 in², or about 2,300 mm²).

In other embodiments, apertures 31 are formed only in inner case 21. Inthese embodiments, apertures 31 are sometimes formed in a single row ofat least one hundred (100) holes, with a total area of at least two andone half square inches (2.5 in², or about 1,600 mm²). In one suchembodiment, apertures 31 comprise approximately one hundred eighty-three(183) holes, with a diameter of about five thirty-seconds of an inch (5/32″, or about 4.0 mm) each, and a total area of about three and a halfsquare inches (3.5 in², or about 2,300 mm²).

In additional embodiments, apertures 31 vary in number and area, and areformed into different numbers of rows in one or both of LPT case 20 andinner case 21. In embodiments with multiple rows of apertures 31, therows are either axially aligned or are staggered (or “clocked”), suchthat individual apertures 31 are rotated about axial centerline C_(L)with respect to one another. Alternatively, apertures 31 are arrangedinto other geometric patterns, or are randomly or arbitrarilydistributed along LPT case 20 and inner case 21.

Like the other components of TEC assembly 10, splash plate (orimpingement baffle) 32 is manufactured from durable heat-resistantmaterials such as metal alloys and superalloys, and is sometimesprovided with one or more protective coatings. Impingement holes 33 areformed in impingement baffle 32 via a combination of drilling,machining, stamping, molding and related processes, as described abovewith respect to main jet cooling fluid apertures 31. Baffle 32 is spacedfrom FODR 12 via a number of dimples or protrusions formed onto baffle32, or via similar spacing means.

FODR plenum 19 extends between a radially inner boundary along FODR 12(that is, along the hot exhaust gas flowpath between FODR 12 and FIDR14), and a radially outer boundary along LPT case 20 and inner case 21(along the cool bypass flowpath inside bypass duct 23). Impingementbaffle 32 extends coaxially within FODR plenum 19, from a forward endlocated proximate finger seal 13 to an aft end located downstream ofprobe assembly 22. Typically, baffle 32 is mechanically attached at theforward and aft ends using bolts, screws, welds or other means ofmechanical attachment, depending upon the particular configuration ofFODR plenum 19 and TEC assembly 10.

Probe assembly 22 comprises probe 34, extending through probe opening35, and a probe boss with various components such as sliding seal orring seal 36 and seal tab or retention ring 37. The probe bosscomponents form a mechanical attachment and pressure seal between probe34 and FODR 12. Probe 34 comprises a sensor or probe device such as anexhaust gas temperature (EGT) probe or EGT sensor, a pressure probe orpressure sensor, a velocity probe or velocity sensor, or another sensingdevice used to measure hot gas path parameters proximate FODR 12.

In typical embodiments, a number of probe assemblies 22 arecircumferentially arranged around FODR 12 and FODR plenum 19. In theseembodiments, individual probe assemblies 22 provide a range of relatedsensing and measurement functionalities, depending upon the desired gaspath parameter measurements. The particular configuration of probeassemblies 22 also depends upon whether TEC assembly 10 is undergoingtesting or is installed on an aircraft for flight operation.

High-pressure (direct-flow) apertures 31 supply cooling fluid to FODRplenum 19 from upstream bypass flow A in bypass duct 23, via one or bothof LPT case 20 and inner case 21. Apertures 31 provide FODR plenum 19with a positive overpressure, as compared to the hot gas path (coreflow) inside TEC assembly 10. That is, the cooling fluid pressure insideFODR plenum 19 is greater than the hot combustion gas pressure, allalong the core flowpath between FODR 12 and FIDR 14. In particular, thepositive overpressure extends along the axial extent of FODR plenum 19,from the upstream (forward) end of FODR plenum 19 to the downstream(aft) end of FODR plenum 19, past probe assembly 22. This preventsinflow through probe opening 35 or probe boss components such as slidingseal/ring seal 36 and seal tab/retention ring 37, or through filmcooling holes as described below.

In designs that do not provide a positive overpressure along FODR plenum19, the plenum is subject to inflow. That is, if the core flow pressureexceeds the cooling fluid pressure inside FODR plenum 19, hot combustiongases can intrude via probe assembly 22, film cooling holes or otherentry points. Hot gas inflow impairs cooling efficiency, resulting intemperature gradients, differential expansion and deformation. Theseeffects, in turn, generate additional inflow and cooling losses. Inflowcan also impair gas path parameter measurements at probe assembly 22,reducing the ability to precisely and efficiently control engineoperations.

In contrast to previous designs, direct flow apertures 31 maintain apositive overpressure in FODR plenum 19 in order to prevent inflow evenunder extreme operating conditions such as power-lift and STOVLoperations. This increases cooling efficiency along the full extent ofFODR 12, which in turn decreases the effects of differential expansionand deformation, reduces wear and tear, and extends service life. Insome embodiments, improved cooling also extends the safe operating rangeof TEC assembly 10, thus improving engine performance.

FIG. 2 also illustrates the structure of splash plate (impingementbaffle) 32, which provides impingement cooling along FODR 12. Inparticular, cooling fluid from direct-flow apertures 31 is supplied atrelatively high pressure, such that main jets J typically flow frombypass duct 23 to FODR plenum 19 under sonic conditions (that is, with aMach number of approximately one).

In designs without splash plate 32, main jets J would impinge directlyupon FODR 12, producing a highly variable heat transfer coefficient thatis relatively high along the main jet line, but attenuates rapidly awayfrom the point of main jet impact. This would subject FODR 12 tosubstantial temperature gradients and thermal stresses, reducing averageservice life and increasing the risk of infant mortality (that is, therisk of separation or other thermal or mechanical failure, beforereaching service life expectancy).

As shown in FIG. 2, however, main (sonic) jets J impinge on baffle 32,rather than FODR 12. Because baffle 32 is located within FODR plenum 19,and not exposed to hot combustion gas flow, baffle 32 is not subject tothe same effects that occur when main jets J impact directly on FODR 12.Baffle 32 converts main jets J into a number of smaller and moreuniformly distributed impingement jets j, which are directed throughimpingement holes 33 to provide more uniform impingement flow onto FODR12 and other components of TEC assembly 20, such as probe assembly 22.

Splash plate/baffle 32 and impingement holes 33 substantially improvecooling efficiency and uniformity along TEC assembly 10, reducingthermal stresses and differential expansion and improving conductiveheat transfer. This in turn reduces mechanical wear and tear andmechanical stress, increasing service life and reducing the probabilityof infant mortality. These effects are not limited, moreover, to theparticular components directly subject to impingement jets j. Becauseimproved cooling reduces thermal gradients and mechanical stress, italso benefits engine components that are thermally or mechanicallycoupled to FODR 12, including finger seal 13, TEC leading edge piece 15,FIDR 14, LPT case 20 and inner case 21.

As shown in FIG. 2, TEC assembly 10 contrasts with designs that supplyFODR plenum 19 via relatively low-pressure cooling fluid plenum 26, andwith designs that do not provide impingement cooling. TEC assembly 10further contrasts with designs that utilize a higher-pressure coolingfluid supply, but do not utilize a splash plate or impingement baffle tomore uniformly distribute impingement flow onto FODR 12, or whichutilize cooling sleeve assemblies or similar structures to direct maincooling fluid jets away from FODR 12, in order to avoid the largethermal gradients associated with direct-jet impingement cooling.

In this respect, the baffled configuration of FODR plenum 19 providesnot only improved impingement cooling but also allows simpler formationof direct-feed (high-pressure) cooling fluid apertures 31. Inparticular, main jet apertures 31 are configurable as simple holes thatdo not require cooling sleeves. This lowers costs, simplifies enginedesign, and reduces the number of potential failure points. Impingementjets j also provide more effective cooling than bulk swirl flow,reducing or obviating the need for tangentially angled holes, angledcooling sleeve assemblies, and other swirl-inducing components.

Positive overpressure within FODR plenum 19 provides greater flexibilityin the positioning of film cooling holes 41. In designs without positiveoverpressure along the complete axial extent of FODR plenum 19, filmcooling holes 41 are restricted, for example, to the forward end of FODR12, as shown in FIG. 2, in order to avoid inflow. Once a positiveoverpressure is established all along FODR plenum 19, film cooling holes41 are locatable anywhere along FODR 12, including downstream regions,without experiencing inflow.

FIG. 3 is an additional cross-sectional schematic of TEC assembly 10,showing film cooling openings 41 extending along the axial extent ofFODR 12. Film cooling holes 41 are formed as apertures, holes, slots orother openings in FODR 12, in order to transmit cooling fluid fromradially outer (cold) side 42 of FODR 12 to radially inner (hot) side43, which is exposed to hot exhaust gas flow.

Film cooling holes 41 are formed into FODR 12 by mechanical drilling,electron beam drilling, laser drilling, laser percussion drilling,electron discharge machining, or another drilling or machiningtechnique, as described above. Holes 41 are typically formed in a numberof annular rows, in which individual holes are either axially aligned orstaggered (clocked), or exhibit a random or arbitrary distribution, asdescribed above for impingement holes 33, and as determined in order toaddress the particular cooling requirements of FODR 12 and otherspecific components of TEC assembly 10.

Baffle 12 converts main jets J into impingement jets j, as describedabove. In the particular embodiment of FIG. 3, the spacing and size ofimpingement holes 33 each vary along the axial extent of baffle 12 andFODR plenum 19. FIG. 3 further illustrates an embodiment in which baffle12 is spaced from FODR 12 at the forward and aft ends, rather than bydistributing dimples, protrusions or other spacing means along the axialextent of baffle 12, as shown in FIG. 2.

Fluid flow through openings 41 generates a post-impingement cooling filmalong hot side 43 of FODR 12, reducing heat transfer from the combustiongas to TEC assembly 10. The cooling film also increases heat transfer bycarrying heat from hot side 43 of FODR 12 into the core (hot exhaustgas) flow, where it passes downstream.

Post-impingement film cooling openings 41 are formed at angle α withrespect to FODR 12, where angle α is measured from the surface to axis αof each opening 41; that is, from the surface toward the perpendicular.A ninety-degree hole angle (α=90°), for example, corresponds to a filmcooling hole that is perpendicular to the surface, while a smaller holeangle (α≦90°) lies more along the surface.

In some embodiments, angle α is approximately twenty degrees (α≈20°), orbetween about fifteen degrees and about thirty degrees (15°≦α≦30°). Inother embodiments, angle α between about ten degrees and aboutforty-five degrees (10°≦α≦45°). In addition, depending upon thefabrication process of film cooling holes 41, angle α can either bemeasured from hot side (hot surface) 43, as shown in FIG. 3, or fromcold side (cold surface) 42.

Typically, axis α of film cooling openings 41 is coplanar with axialcenterline C_(L), such that angle α is a downstream angle (or“streamline” angle), designed to encourage attached flow in a downstreamaxial direction along hot side 43 of FODR 12. In some embodiments, filmcooling openings 41 also have a circumferential component, as measuredtangentially along hot surface 43 (or cold surface 43), andperpendicularly to central axis C_(L). In these embodiments, thecircumferential angle typically varies between ten and forty-fivedegrees)(10-45°), in order to encourage tangential film flow along hotside (hot surface) 43 of FODR 12.

In previous designs, film cooling holes were limited to the forward endof FODR 12, as shown in FIG. 2, because the cooling fluid did not have apositive overpressure all along FODR plenum 19. This was particularlyproblematic at the aft end of FODR 12, where the hot gas path pressureincreases due to expansion, stagnation and recirculation, leading toinflow as described above. Where inflow occurs, film cooling isessentially impossible, or at least substantially degraded in bothefficacy and efficiency, because the holes must be located upstream ofthe desired cooling location.

With direct-flow (high-pressure) cooling fluid apertures 31, FODR plenum19 maintains a positive cooling fluid overpressure all along the axialextent of FODR 12. This prevents inflow and allows film cooling holes 41to be located along the complete axial extent of FODR 12, including aftregions downstream of probe assembly 22. This allows film cooling holes41 to be located and sized on the basis of cooling needs, rather thanbeing limited by the core gas path pressure.

The result is greater cooling efficacy and efficiency. Efficacy improvesbecause film cooling holes 41 are located where most needed, includingaft (downstream) locations along FODR 12, and are not restricted toforward (upstream) locations. Efficiency also improves, because lesscooling fluid is required when film cooling holes 41 are located wherecooling is required, rather than being placed upstream (forward) due topressure considerations.

Positive plenum overpressure also allows cooling holes 41 to be moreuniformly distributed along the entire axial extent of FODR 12,decreasing temperature gradients and thermal stress, and reducingdifferential expansion and mechanical stress. This also has beneficialeffects both for FODR 12 and the other components of TEC assembly 10that are thermally or mechanically coupled to FODR 12, as describedabove for impingement baffle 32 and impingement holes 33.

The present invention has been described with reference to preferredembodiments. The terminology used is for the purposes of description,not limitation, and workers skilled in the art will recognize thatchanges may be made in form and detail without departing from the spiritand scope of the invention.

The invention claimed is:
 1. A system comprising: a bypass duct having aradially inner margin and a radially outer margin; a turbine case formedalong the radially inner margin of the bypass duct; a flow ringcoaxially disposed within the turbine case; a splash plate extendingaxially along the flow ring, wherein the splash plate is radially spacedbetween the flow ring and the turbine case; a plurality of cooling fluidapertures formed in the turbine case to provide cooling fluid flow fromthe bypass duct onto the splash plate; a plurality of impingement holesformed in the splash plate to provide impingement flow onto the flowring; and a plurality of variable area vanes positioned downstream ofthe cooling fluid apertures in the bypass duct.
 2. The system of claim1, wherein in use the cooling fluid apertures provide substantiallysonic cooling fluid flow onto the splash plate, the cooling fluid flowhaving a Mach number of approximately one.
 3. The system of claim 1,wherein in use the cooling fluid apertures maintain a positive coolingfluid overpressure as compared to hot combustion gas flow in the flowring.
 4. The system of claim 3, further comprising a plurality of filmcooling holes formed in an axially aft portion of the flow ring.
 5. Thesystem of claim 4, wherein the film cooling holes are formed along asubstantially full axial extent of the flow ring.
 6. The system of claim5, wherein the plurality of cooling fluid apertures comprises ten ormore apertures with a diameter of at least one half inch (at least 12.7mm).
 7. The system of claim 5, wherein the plurality of cooling fluidapertures comprises one hundred or more apertures with a diameter of atleast five thirty-seconds of an inch (at least 4.0 mm).
 8. The system ofclaim 1, wherein the variable area vanes are configured to increase apositive overpressure of the cooling fluid to prevent hot gas inflowalong a full axial extent of the flow ring.
 9. A cooling system for aturbine assembly, the system comprising: a turbine exhaust flowpath; abypass flowpath coaxially oriented about the turbine exhaust flowpath; acooling plenum coaxially oriented between the turbine exhaust flowpathand the bypass flowpath; a plurality of jet apertures to provide coolingfluid flow from the bypass flowpath into the cooling plenum, wherein thecooling jet apertures maintain a positive overpressure along a fullaxial extent of the cooling plenum; a baffle extending axially along thecooling plenum, the baffle spaced radially between the turbine exhaustflowpath and the bypass flowpath and having a plurality of impingementholes to provide impingement flow from the cooling plenum onto theturbine exhaust flowpath; and a probe assembly extending from thecooling plenum into the turbine exhaust flowpath.
 10. The cooling systemof claim 9, further comprising a plurality of film cooling holes toprovide cooling fluid flow from the cooling plenum to the turbineexhaust flowpath.
 11. The cooling system of claim 10, wherein the filmcooling holes extend axially downstream of the probe assembly.
 12. Thecooling system of claim 9, further comprising a plurality of variablearea vanes positioned in the bypass duct to increase the positiveoverpressure and prevent hot gas inflow from the turbine exhaustflowpath during short takeoff or vertical landing operation of theturbine exhaust assembly.
 13. A turbine exhaust assembly comprising:inner and outer turbine rings coaxially disposed about an axis, theinner and outer turbine rings forming an exhaust gas flowpaththerebetween; a turbine case coaxially disposed about the outer turbinering, the turbine case and the outer turbine ring forming a plenumsection therebetween; a splash plate extending axially along the plenumsection and spaced radially inward from the turbine case, the splashplate having a plurality of impingement apertures for impingement flowonto the outer turbine ring; a plurality of jet apertures formed in theturbine case, the jet apertures configured to direct bypass air into theplenum section with sufficient pressure differential to prevent inflowfrom the exhaust gas flowpath along a full axial extent of the plenumsection; a bypass duct coaxially disposed about the turbine case; and aplurality of variable area vanes circumferentially arranged about thebypass duct, downstream of the jet apertures.
 14. The assembly of claim13, further comprising a plurality of film cooling holes formed in theouter turbine ring to provide cooling fluid flow from the plenum sectioninto the exhaust gas flowpath.
 15. The assembly of claim 13, wherein thefilm cooling holes extend substantially along the full axial extent ofthe plenum section.